text:
What is the heaviest payload a modern theoretical glider could carry?
Let's define a glider as any plane with a better than 0.75m/s minimum sink rate. Let's also assume it meets all EASA requirements for load factors and handling.
Knowing that the spar is an important limiting factor-- if not the limiting factor--  for payload capacity, and that the spar capacity is a function of material constraints, how much could a theoretical modern glider carry before the wingspan simply couldn't be increased any more?
I'm looking for answers to within a power of 2.

Hard to say but I'd guess in the hundreds of kg without reaching 500.

Why could you not design a glider with the wingspan of a 747 (but better low-speed aerodynamics)?  Or even larger, since wingspan of commercial planes is constrained by the need to park at gates?  You could even go with a multi-hull design, like this: https://en.wikipedia.org/wiki/Scaled_Composites_White_Knight_Two

The sink rate appears to be the restriction to take into consideration. Otherwise, I would say, any aircraft could be viewed as a "self launching glider" so the payload would be quite large.

Some were built to carry 20 metric tons or 200 pax with bag and baggage.


text:
Looking for published version of a successful aircraft certification effort
I'm assuming-- and please correct me if I'm wrong-- that at the end of the certification process, the FAA receives a thick binder of materials. Is there an online link to an example of said binder?

It doesn't really matter how old it is, I'm just looking for something representative.
If beggars can be choosers, simpler is better: I'd rather peruse a binder for a J-3 than for an A-380.
The certification authority is not important, e.g. EASA and FAA would be equally good.

Not sure the info would be public, there is potentially a lot of proprietary design details in the package...

@MichaelHall I've been wondering about that. I've no idea if design detail confidentiality could/should/would survive an FOIA request. Presumably, it is in the public good to be able to independently verify that certification standards are being met. That being said, I don't want to bloviate on what could be.

It's more like a endless stream of reports.  Qualification test reports, Endurance test reports, Structural test reports, System safety analysis reports, Flight test reports and on ad infinitum. On a transport airplane the cert package will fill a room.  On a light plane, maybe just a bookshelf.

Good point by John K.  I've only been involved in certifying electronic components, but one part could fill a few binders.  And to my earlier remark, the test plans and reports tend to not have much design details, more performance requirements and qualification test levels achieved.

@MichaelHall "performance requirements and qualification test levels achieved" <-- Those are exactly the kinds of reports I'm interested in reading. I'd like to see how the reports look and feel, how extensive the proofs are, etc... The rabbit-hole of "Yeah, but did you really prove it? What about the test machine's calibration? What about the calibration machine's calibration? What about..." is potentially infinite, so I'm interested in what best practices look like.

source: Stack Exchange

license: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/4.0/

sort: votes


text:
When are "winter months" when referring to runway closure?
1B9's grass strip is "CLSD WINTER MONTHS", https://www.airnav.com/airport/1B9.
What is the definition of "winter months"? Is it formally Dec 21st through Mar 21st? Or is it just when it's cold outside, for whatever the pilot decides is "cold"?
If it's not NOTAM'ed as closed, is it therefore legally open?

This seems like a simple "common sense" situation to me. It is a turf strip. It is closed "winter months and after heavy rains." If there is any possibility of there being snow on the ground, or that the turf is waterlogged, you shouldn't land on it...

My secret decoder ring says that means “we don’t have a snowplow.”

It's just a legal butt-cover statement put there by the airport owner to absolve the owner of any liability if someone prangs their plane after it snows, or tries to land when it's soggy and flips over in the mud, and tries to blame the airport operator.
It's a warning that there is no winter maintenance and the place maybe isn't that well drained in spring or after a thunderstorm, so you're on your own.  No calendar date is implied and there is no legal implication from a regulatory standpoint.

Well, it could also have to do with owner not desiring ruts to be left in the turf!

all_licenses: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/4.0/


text:
Are there instances of ATC services being refused to targeted pilots and/or operators?
Jpe61 opines that in certain circumstances aircraft can be denied services because of overdue fees. One service mentioned is ATC. This makes me curious if this could actually happen, since for certain airspaces ATC is federally required. (For ex., can't take off from a Class B without clearance.)
Can federally-mandated ATC services be denied in isolation (i.e. not as a result of revoking a license or permit), and, if so, are there instances of this occurring?

I think you should ask to the law stackexchange. And possibly you should carefully read the contract and term of service of the fuel operator. FAA services cannot help you to bypass existing laws (but if FAA is authorized explicitly to do so).

The question you linked concerns the UK airline Thomas Cook and EU/UK laws. To my understanding users there pay a fee for ATC services. This is not the case in the USA; over here ATC is funded primarily by taxes on airline tickets and taxes on aviation fuel, rather than a specific fee paid by the pilots/operators directly.

I don’t know how this would even be possible in the US system.  I fly out of a tower-controlled airport (that underlies a class B).  When I call and say “N12345 is ready taxi with Alpha, departing VFR northwest” they aren’t looking me up in a database, checking my registration, etc.  I could even have given a wrong tail number, or have an expired registration, or even an invalid tail number (this happened to me personally) and no one in ATC will know or care.


text:
Are maintenance manual airframe lifetime limitations binding?
An IA told me that component lifetime limitations, even when published in the Airworthiness Limitations section of the service manual, are not binding. They can be maintained on condition, completely superseding the published instructions.
In a related vein, I found the following section in an airframe manual:

Extension of airframe life exceeding 5,000h can be done for individual aircraft only after a comprehensive airframe inspection carried out by the manufacturer according to an approved program.

Under what conditions does on condition apply to published airframe lifetime limitations as well?
Does it matter if the manufacturer is still in business?

If the manufacturer is not in business you will have to talk to the local FSDO to find out if there is a way to comply with the airframe inspection. Your IA may be correct from a FAR point of view (and that is debatable), but an insurance company will likely deny any claims for any incidents or accidents if you did not comply with the manufacturer's maintenance and inspection instructions, and a court would likely side with them, leaving you holding the bag for any damages or liability.

Thanks, that's helpful to know. Do you have links to flesh out why the IA's opinion may be debatable, as well as why the insurance and court will likely deny any claims? That's a very powerful assertion, and while it has the ring of truth to it it would be nice to see that backed up by case law and/or opposing legal interpretations.

In a battle between an IA and a FSDO, the FSDO almost always wins. Ask the IA if he/she will put his/her opinion in writing and defend it against the FSDO saying otherwise. That will be your sign. As to what happens with your insurance claim, call any aviation insurance broker and ask. If I were you, I would find out who owns the type certificate for the aircraft and go to them. Links are helpful but common sense more so. Completely abandoned type certificates are few and far between. Which aircraft is your question about?


text:
Would reverse thrust increase with a decrease in aircraft forward speed when referring to propeller reverse thrust?
Would reverse thrust in a propeller increase with a decrease in aircraft forward speed?

If the propeller keeps the pitch constant while in reverse thrust,  the component of the relative wind perpendicular to the prop disk goes down with decreasing aircraft speed. Hence, the AoA also goes down, the lift produced by the blade decreases, and the reverse thrust does also decrease...

That assumes that the engine RPM is also constant. Reducing the load will usually increase the RPM, and higher RPM could produce more power. So "it depends"


text:
What is the purpose of the extending wingtip panels on the Su-25?
Is it to increase the air resistance of the aircraft?

Source: reddit.com

Those are airbrakes, they allow the aircraft to increase its drag substantially to bleed off excess energy. They are hydraulically operated, and need pressure for both deployment and retraction, i.e. they do not auto-stow, at least on this aircraft.
Airbrakes are a handy feature, especially for a ground attack aircraft like the Su-25 in your picture, as they help both with landing on short airstrips and with dive bombing, a tactic that requires to dive almost vertically on a target to minimize slant errors in targeting. This is made possible by the airbrake, which prevents the aircraft from exceeding its dive speed.
Other aircraft types use or have used airbrakes in the past, mainly to dissipate energy during an approach. Examples include some airliners, most fighters and nearly all gliders.
Fighters are usually optimized for high speeds which are dangerous to land at, and an airbrake enables them to quickly slow down while on final approach, remaining at a comfortable speed until the last moment. They also have some occasional uses in WVR (Within visual Range) combat, but this will be rabidly argued against by the BVR (Beyond Visual Range) and stealth cults.
Gliders also use airbrakes to control their speed on approach and when diving, but also to arrest undesired climbs, when caught in a thermal when they need to descend. Since typically gliders lack a hydraulic or even electric system, their airbrakes are actuated by the pilot's muscles and are more akin to spoilers.
The geometry of an airbrake can vary significantly, but the principle is the same: a large surface extended into the airstream that generates mostly drag. their location on the airframe is a compromise between structural, mechanical and control considerations, i.e., it needs to be strong enough to resist the loads, easy enough to route hydraulic lines or control rods to, and not significantly impact the aircraft's stability.  The examples below are a BAe 146, an F-15, a Slingsby Capstan and a MiG-23:

I'd love to know why the Su-25 has such unique speedbrakes though. They appear to be parallel to the direction of travel, which seems odd for a brake.

@zymhan It´s hard to find a good high-res picture online, but if you trust Trumpeter models to be accurate, you can see the geometry here; it is not parallel to the airflow at all. If I had to wager a guess about why it is there, it is likely the designers did not find enough space around the tailcone to deploy them out of the engine exhaust plume.

Late answer but.....extending speed brakes/airbrakes during an approach allows the aeroplane to fly slower because it changes the aerodynamic characteristics (specifically the relationship between lift and drag).  The drag caused by the speedbrakes increased the total drag but reduces the speed where the "drag curve" graph is at its lowest.
As a bonus, the extra total drag requires a higher power setting, so if the pilot has to break off the approach (go around) the engines are already running at higher revs and so take less time to spool up to full power.  Retracting the speedbrakes at that point reduces the total drag and together with rapid engine spool up, helps the aeroplane to accelerate more quickly

I'm having a hard time believing these airbrakes allow for slower flight. They do not change wing profile, so they will not change lift curve. With or without the airbrakes in question, the wing will stall at the same AoA, and the same airspeed.

While more parasite drag may shift the Min L/D point to the left (i.e. slightly slower), an approach is typically flown at a margin above stall speed, rather than simply at min L/D. The second paragraph is good info, but I'm also doubtful about the first paragraph's answer.

Jpe61 Lift-induced drag remains essentially unchanged but when plotted on a graph shows a curve that Decreases as speed increases.  Form/parasite drag increases with the extension of the speed brakes - and when plotted shows a curve that Increases with speed.  When you plot the two lines on the drag Vs speed graph, the total amount of drag increases but the point where they intersect (minimum drag speed) occurs at a lower Indicated air speed.  The net result is as I explained earlier.

As a former instructor on RAF Harriers and Tornadoes, as well as being now an instructor training airline pilots, I can promise you that what I have written is 100% correct :)  Trust me  I've been there.  To be frank, the final stages of flap on an airliner have a broadly similar effect - contributing little extra lift but lots of extra drag, shifting the drag curve up and left (to a higher total drag and lower speed)


text:
How to calculate the number of blades?
Inspired by building a new drone, I got lost in curiosity.
What if I have a fixed length of rotor blades and a target amount of lift I want to, well, lift. How do you calculate the number of blades?
I have read a lot about disk loading and was intrigued by this answer (https://aviation.stackexchange.com/a/32252/44213) which found a correlation between diskloading and amount of blades (or lift per blade). It made me wonder; is it actually possible to calculate the number of blades? If it matters (and I guess it does) I want to use a naca0012 airfoil.
Searching around I did find the relation between the number of blades and decreasing efficiency/lift per blade when you add a blade. Also the fact that is would be easier to lengthen the blades is something I found. But if lengthening is not an option, how can you actually calculate this? And, to complete the question disturbing me, what effect does a second disk (like the chinook) have on these calculations?
Any help is much appreciated. Any links will be read thoroughly.

Determining the number of blades is one of the steps in designing the main rotor of a helicopter, but not the first one. The main rotor design process is an iterative process with quite a few input parameters, which depend on the mission type:

Max. TO weight
Max. speed
Installed engine power
Autorotation performance
Noise
Manoeuvring g
Blade profile
Blade area
Vibratory loads

Using these inputs, the design process determines:

Rotor diameter. This determines the disk loading: lower disk loading results in better hover performance and auto-rotational performance, higher disk loading in better cruising flight performance.
Tip speed. As close as possible to the drag divergence speed in cruise, also a limiting factor in selecting rotor diameter.
Rotor solidity. How much disk area is actually occupied by blade area, determined by blade stall limits in manoeuvres. Increasing the rotor solidity requires more engine power, but increases the blade stall margin.
Number of blades. Mostly based on dynamic criteria, less on aerodynamic ones. The Bell 212 shows that rotor solidity can be increased by using wider blades, but using more blades minimises vibratory loads and the noise per blade passing.

The rotor design process is quite broad, and can best be found in the theory books which also contain statistical data to aid the initial stages. The plot below is from Leishman and shows that light helicopters have a blade length of about 4 meters, apparently a good length for this weight and mission type.

The overlapping twin rotor disks of the Chinook result in one rotor operating in the wake of the other, reducing efficiency and requiring more disk area. So total disk loading will reduce.
Two good books on helicopter stuff: Helicopter Aerodynamics, Stability and Control by Ray Prouty, and Principles of Helicopter Aerodynamics by J. Gordon Leishman.

Thanks so far. What I try to understand is, what does an extra rotorblade actually add in lift, percentage wise? And what should be deducted for the overlapping disk area at e.g. the ch47? It surprised me that the diskloading/m2/blade for many helicopters is about 8.7, but for the ch47 it is 15,3. Do you have some gestures to help me understand?
BTW: I start thinking from scratch, so the other parameters can be dependent variables.

“... what does an extra rotorblade actually add in lift, percentage wise?” Depends on what we are comparing, add a rotor blade to a rotor that can already lift a helicopter means we’re only adding drag if we don’t keep total blade solidity equal. The disk loading per blade is not a main initial aerodynamic design parameter and we cannot look at only that without considering all other related parameters. The answer you link to suggests that blade loading is a main design parameter, which is not really how it works.

OK, I think I understand what you mean; normally the blade is designed to fit the needs of the rest. But, what if you want to design a drone and have to work with available blades. Not to be picky, I just like to understand the matter and sometimes I end up thinking the other way around than it is normally done. It helps me to understand why. Sorry for that.

Some calculations here


text:
What is this quad-jet plane at El Paso Airport?
If you look at El Paso International Airport on Google Maps, on the south side apron and across from a hangar that is labeled "Sierra West Airlines" (west of the NASA facilities) there is a small quad-jet plane:
The aircraft is about 20 m long and has a wing span of about 18 m:

Any idea on what this is?

I added an image from Google maps, I guess this is the one you were talking about?

This appears to be a McDonnell 119/220, a 1950s vintage business jet attempt (from before McDonnell merged with Douglas).  The four engine layout with engines under the wings, coupled with the size, is very distinctive.
The one at El Paso is apparently a genuine rarity — there was only a single unit (the prototype) ever constructed.

It's the McDonnell 220, registration N4AZ.

(Source)
Wiki says

The McDonnell 119/220 was an attempt at making a business jet by McDonnell Aircraft in the late-1950s. It had a configuration that was unique by bizjet standards, with four jet engines mounted in individual pods underneath a low wing; it could accommodate ten passengers in a luxury executive configuration but could carry as many as 26.
...
The McDonnell Corporation used the airplane as a VIP transport for a few years before donating it to the Flight Safety Foundation's research facility in Phoenix, Arizona. As of 2020 the single prototype is still extant, registered as N4AZ and stored at El Paso International Airport.

Only one such plane exists. Its first flight was in 1959.
Here are some pictures.

good god that had to guzzle some gas, although not much different than a Jetstar. Thanks for the answer.

'Dale King'  was impressed that you and Zeiss Ikon (25 seconds after you :-) )  were able to produce that answer so well and rapidly. He said:
"Signed up just to say how next level the internet is. You can just post a picture of a pixelated plane and within an hour someone tells you what it is." What a time to be alive.

@WilliamTaft According to the Wikipedia article we both linked, it appears that it was in direct competition with the Jetstar for a military light transport contract (the kind now held by VC-20).  It was 1959, so anything with jet engines needed big tanks to have range...

@RussellMcMahon I was similarly impressed by the answers to a question I asked last summer. The user DeepSpace could within short time identify and produce the flight path of the plane I had observed. I know that's commonplace for you guys;  but to me as an aviation layman but computer guy it is an indication for the scope of the information age, the level of transparency we are heading towards  and the corresponding disappearance of privacy.


text:
How do slats reduce stall speed?
I've read in PHAK that slats reduce stall speed, but in what way? Is it because of the increase in surface area?

Could you include a link (or better description) to what PHAK exactly is?

I'm sorry, but I actually read it somewhere else (I forgot where, I read too many stuff), but PHAK is Pilot's Handbook of Aeronautical Knowledge by FAA ( https://www.faa.gov/regulations_policies/handbooks_manuals/aviation/phak/media/pilot_handbook.pdf ).

In some very rare cases they don't just reduce stall speed but eliminate it altogether.

Nice info! But it says there that it only has a very low stall speed (50kph), not necessarily eliminated?

It's because leading edge devices allow a higher Angle of Attack.

The four types of leading edge devices work by pointing the nose of the wing downwards so that at higher AoA there is no flow separation at the upper surface near the nose. Deflecting the slats does not increase $C_L$, the aeroplane must increase AoA to do that. Lower stall speed comes with the higher $C_{Lmax}$

From this answer, which also contains a graph showing that flaps increase $C_L$ at constant AoA.

The leading edge flaps are the ones which deflects downward, and the slats technically extend forward to create slots (which prevents immediate airflow separation) don't they? I'm a bit confused with the main difference of these two, because they function the same

Yeah their function is the same, of all four: re-align the nose to be more in line with the free stream at high AoA. Note that if increased wing area would be a factor, the $C_L$ would rise at constant AoA, which it doesn't.

The two main leading edge devices are leading edge slots and leading edge flaps:
Slots energize the boundary layer to delay stall by allowing higher pressure air from below to "leak" in a controlled way to the upper surface.  The slot is effectively a convergent duct, large at the bottom and small at the top, so the air is slightly accelerated as it flows through the outlet at the top, more or less parallel to the skin surface.  Because the air is slightly accelerated relative to the freestream just above it, being squeezed through a convergent duct in effect, Coanda effect enhances its ability to follow the upper surface contour and help the freestream stay attached.
Leading edge flaps, basically drooping leading edges, increase wing camber the same way that a trailing edge flap does, increasing Clmax somewhat at a given angle of attack.
If I combine the leading edge SLoT and flAp together, I get a SLAT.  The slat extends its nose down and forward from the normal LE position, increasing camber somewhat, like a LE flap.  The slot created when the slat extends energizes the boundary layer and delays separation, allowing the wing to generate lift into the mid 20 degree range.  So you get a little bit higher CLmax for a given AOA and it can operate to a much higher AOA, giving you the highest possible CLmax and lowest stall speed.

Fixed slots were all the rage in the 30s and 40s to control flow separation forward of the ailerons on airplanes like the Globe Swift, negating the need for wing washout, but these didn't really reduce stall speed, just kept the ailerons working in the stall.  They fell out of favour because the drag penalty of the always open slot was worse then the efficiency loss of washout.

I see what you did there. Very nice way to remember it!

Slats increase the camber of the wing, which increases the coefficient of lift.
When you deploy slats, the AOA is actually lowered for that part of the wing.  This is the same washout principle Dunne ingeniously applied to the D.8 swept wing bi-plane in 1912!
You now have a more heavily cambered wing at a lower angle of attack, so more lift is to be had by increasing pitch more.  Since the Clift is higher, one can generate the same amount of
lift at a lower speed.
Airflow underneath the wing and its contribution to lift may be worth further study.  The action of the air curling around the wing leading edge and impacting the "underside" in an upwards and forward manner (Eye of Jupiter?) or a bit of compression may help explain why they work so well (in addition to the benefits of lift creation by increasing camber on the upper part of the wing).
Slats are greatly helpful in reducing the AOA and increasing the Coefficient of Lift of part, or all, of the wing. You can now pitch up a little more, but be careful.

Slats are greatly helpful in reducing the AOA and increasing the Coefficient of Lift is not what they do. Leading edge devices don't raise $C_L$ at constant AoA, see graph. They are not helpful in reducing AoA, they allow for greater AoA without stalling.

Mate - please look it up. AoA as defined by the angle relative to the free stream.

But it’s OK, I can delete the comments if you like.

Right, look at the chord line angle to the free stream after the slat is dropped.  What is there is a more cambered wing at a lower AOA.  So that PART of the wing now has different properties.  Lower AOA AND more camber. So you've washed out the wing tip (relative to the rest of the wing). The reason why this is so important (slats really helped the Sabre Jet (the first of the too small horizontal stabilizer (guess what happens with a swept wing))aircraft) to understand  their role in lowering AOA relative to the rest of the wing and assisting horizontal stabilizer in preventing stall.

Because when an aircraft goes to excessive AOA, the lift of the horizontal stabilizer should be going positive to help lower AOA, even without neat tricks like pushing the stick forward (ha, ha, ha).  Good slats will also help as will generate a pitch down as they (on a swept wing) are behind the CG, and by virtue of their increased lift at higher alpha (from your graph).

Yes that is my point, you’re talking about the local AoA relative to the nose, Best specify that.


text:
What is the drag optimized wing form for given lift and bending moment?
I am trying to build a passive hydrofoil board.
Something like this:

I want to use wood and i am coming to the conclusion that i want to optimize drag for a given lift and bending moment.
Doing my research i found out that an elliptical wing isnt the best approach.
But a linear tapered wing should work better.
Is there a rough formula for this optimization?
I already came across two articles:
"Lift Distributions for Minimum Induced Drag with Generalized Bending Moment Constraints" and
"THE SPANWISE DISTRIBUTION OF LIFT FOR MINIMUM
INDUCED DRAG OF WINGS HAVING A GIVEN LIFT AND
A GIVEN BENDING MOMENT"
I only need a rough formula/shape

You have already found the most relevant articles. In 1933 Ludwig Prandtl published "Über Tragflügel kleinsten induzierten Widerstandes" in the Zeitschrift für Flugtechnik und Motorluftschiffahrt. 17 years later R.T. Jones came to the same conclusions in NACA-TN-2249. The 2013 article "Lift Distributions for Minimum Induced Drag with Generalized Bending Moment Constraints" by David J. Pate and Brian J. German goes one step further and optimises drag at a lower and structural mass at a higher lift coefficient.
But that is not all that relevant to your case. You want a wing which supports your weight plus the small surfboard which sticks out of the water. At those small loads, structural mass is almost negligible. Scaling laws mean that structure becomes more important as aircraft grow in size and mass. You know that elephants have much more massive legs relative to their body size than antelopes (or even ants, for an even more drastic comparison), since body mass scales with the cube of linear dimension while structural strength scales only with the square of linear dimension. This means that wing spar weight will be proportionally higher for larger aircraft.
As a consequence, insects have more elliptic wings than albatrosses, and model aircraft have optimum wings which are much more elliptic than the optimum wing of an airliner. The optimum shifts from an elliptic load distribution at very small scales to an almost triangular distribution at large scales.
What is more relevant is the angle of attack range in which this wing is meant to operate. The pumping motion seems to lead to quite large angle of attack variations, so you cannot optimise that wing for just one flow condition. Instead, it must be robust enough to produce tolerably low drag over the whole operating range.
Drag has two components: Pressure drag and friction drag. Pressure drag has one part which stems from lift creation, which in turn depends on the mass of fluid affected by the wing. That part drops with speed and wing span, so you want to be fast with super wide wings in order to minimise that kind of drag. However, friction drag increases with speed, so there is an optimum which varies with wing planform.
But there is another issue. Wide wings are very sensitive to angle of attack variations while more narrow, stubby wings will tolerate a wider angle of attack range. Also, the increased chord length of more stubby wings produces less friction drag per area than the narrow but wide wing.
Next to consider is the relative thickness of the wing's cross section (called an airfoil) and its camber. A thicker airfoil will create more static lift (buoyancy) but will create less dynamic lift. Make the airfoil too thin and its useable angle of attack range will become very small. I would select a thickness between 12% and 15% of chord length. Camber should reflect the operating lift coefficient: For highly loaded wings, use more camber. The tail surface, however, should be uncambered.
Another important factor is the incidence of wing and tail. The tail should have less for positive static stability, but you need to include the wing's camber into that calculation.
All those questions can only be answered when we know the flow speed, the mass on the board and the size of both surfaces. But you have to start somewhere. For that, it might be best to start with a proven design, collect the operating parameters and then go back to the drawing board with a sound basis for optimization.
For a first shot, start with a trapezoidal wing planform in which the tips have 70% of the chord of the root. Aspect ratio should be 5 or 6. You may add highly swept tips and 3D-shaping as you desire, but don't overdo it.

Thank you so much. I already do have a first outline. I am trying to optimise it for 3m/s. The wood i am using now can handle the bending moment xflr calculated for my foil. I feel like i lost the overview though. To reach my goal i have to optimise poweroutput. i want to be able to start from a deck. exactly like in this video.


text:
How do boxwing and tandem airplanes balance moment
I want to build a hydrofoil like this picture:

I discovered that you can pump it with a leg motion and moving the center of mass.
Like this:

I discovered xflr5 and did some of my own simulation. I am in aww of the complexity of aeronautic engineering. I learned so much in such a small time.
As you can see in the video, pumping is very hard and you can only do it for short amount of times. so that i had the idea to optimize the wing configuration for this type of drive. I am not quite sure if the rear wing produces lift or is actually pushing down like in a commercial aircraft. My design in xflr5 has a very high aspect ration and a very high wingspan and i am searching for designs with higher aerodynamic efficiency.
I found a few ideas:

Flying wings i guess it would be hard to get longitudinal stability if the center of mass is so high.
A canard wing configuration.
Flying wing with bell shaped lift distribution
Boxwing configuration
Tandem configuration.

But i dont understand how tailless boxwing configuration produce longitudinal stability same for tandem configuration.


text:
Why are there no stealth helicopters?
There aren't exactly an abundance of stealth planes, but most highly developed countries own one. Why are there no stealth helicopters? If there are, why are they so obscure?

I'm not sure we can accept the premise that there are none. It's widely speculated that the US used a stealthy helicopter to execute the Bin Laden raid. And the US government has a history of keeping certain aircraft as secret as possible.

There are. But the meaning of stealth is a bit different. For helicopters the emphasis is noise and IR rather than radar, and this is part of the reason why Comanche is canceled.

@ymb1 I guess that was a copy and paste mistake :) I can't fix it now, and I think both questions showed up in the related list automatically anyway. I'll just delete it.

@zymhan The book Relentless Strike goes into a fair bit of detail about that stealth helo. After reading the chapters on that topic in that (excellent) book, you'll dismiss the premise of this question.

thank you! I did read a bit on the RAH 66 Comanche, and a better question would have been, "Why haven't there been any attempts at a stealth helicopter since the Comanche?" And in regards to the Bin Laden raid, I did see those pictures, and I wasn't thinking about that when I asked the question. Thank you.

Well there actually are stealth helicopters, from Boeing Sikorsky’s RAH-66 Comanche, to a classified low observable derivative of the UH-60 Blackhawk, which was used in operation Geronimo to kill Osama bin Laden.  They’re out there but haven’t seen the light of day for several reasons.
UPDATE: The US Army has procured a new scout / light attack helicopter program with the Bell 360 Invictus program under development.  The Invictus, a two seat attack helicopter which incorporates low observable features appears to follow many of the configuration choices made with the RAH-66 nearly two decades ago.
In the case of the RAH-66, It was development problems and massive cost overruns which led the secretary of the Army to cancel that program.    It would have been interesting to see what the production version of the Comanche had in terms of an RCS, as with the electronics it carried plus the integration of the brimstone anti-tank missile would have made it very formidable.
It’s possible as well that with the Comanche and the classified Blackhawks that the military has had a chance to evaluate these systems and determine that low observable applications and helicopters was either not beneficial for their domains of operation or simply superfluous.  In the case of the classified Blackhawks, those items are sort of delta force type weapon system used for special, clandestine missions and there may not be a large need for a fleet of these things.
Given most military helicopters are operating nap of the earth, and the major threat to them comes in the form of small arms fire at visual distances, low observable technology (at least for radar) really isn’t much help in this situation.   Helicopters do make use of some low observable technology, particularly IR suppression systems for the engine exhaust, to make them less vulnerable to MANPADS, but radar detection really isn’t a big threat to them, as they generally operate well under the observable ranges of AAA and SAMs used in IADS.  A Zeus, however, is a very dangerous thing to face down in a helo; it’s possible that LOT may be of help here.

I looked up what a Zeus is: it's a ZSU-23-4 Shilka, a Soviet armored, tracked anti-aircraft weapon system with four 23 mm gun barrels introduced in 1965 and still in use today.

The closest question to this currently is Why aren't modern helicopters quieter?, which one should read first. That regards sonic stealth.
As for radar stealth, any rotating part, visible from the outside, is a major source of unavoidable radar reflections. In 4th generation jet fighters, the engine faces generally contribute more than half of the total RCS (radar cross-sections). Very simple tweaks like installing radar blockers into the intakes can reduce the RCS several-fold.
The rotor is a very large rotating part. Even if made completely radar-transparent, the rotor will accumulate dirt and grime from the air, which have some reflectivity. On a properly stealth-shaped surface, this will be deflected away from the mainlobe. But a rotor, constantly rotating and flexing, has an uncontrollable shape.
This means helicopters can never reach B-2 or F-22 level of stealth. They can get lesser RCS reduction, along the lines of the Silent Eagle proposal. This could be useful, but helicopters fight at much shorter ranges than fighters or bombers.
The F-15 is expected to be visible to high-end opposition fighters out to 300-350 km, and the F-22's stealth reduces that to about 40-50 km against most fighters. At that range, radar stealth is defeated by more powerful or longer-wavelength radars, and even an aircraft with no radar cross-section could be detected by IRST/FLIR systems (correspondingly, high-end opposition fighters use all of these). The F-22's weapons have some hit probability at ranges up to ~80 km against fighters and up to twice that against other aircraft, creating useful room to shoot first. Helicopter radars top out around 50 km and their weapons even closer.
The Comanche could never hope to be as stealthy as the F-22. It's hard to guess its best-case scenario, but something like 60-100 km would've been pretty good. Operationally, for a helicopter, such detection range doesn't make much difference.
Helicopters can already hide from THAAD (theater area air defenses, mostly networked long-range SAM such as the Patriot or S-400) by nap-of-the-earth flying, placing surface features between them and the radar, and hiding among the ground clutter. It's only at 15-40 km that a helicopter would come in direct view of short-range air defenses (highly mobile "all in one" transporter, launcher and radar vehicles). A Comanche would still be visible and targetable by radar at this range.
Essentially, a helicopter's stealth would help protect it from enemy fighters. But there's more than one way to protect your aircraft from enemy fighters.
For the US, it's air superiority. The US maintains an air force that, including the USN, is a match for the rest of the world combined for air-to-air combat. It's not by happenstance - US military doctrine prioritizes air superiority, so resources were spent on it.
With air superiority, a helicopter doesn't need to hide from enemy fighters; the risk they face in every mission won't be worth it for such a low-priority target. The Comanche project was started in a wave of excitement with stealth, with an assumption that stealth could be an easy add-on to it. Without stealth considered, it was underwhelming. Radar-stealthing a helicopter proved difficult and expensive, and there just wasn't much to be gained by it.

Also, the F-22 can cover that 40-50 km a lot faster than any helicopter. Even if a stealth attack helicopter can be sighted by a defence radar at the 30km range, travelling at 300km/h will give the defence entire six minutes to prepare a response

"The rotor is a very large rotating part. Even if made completely radar-transparent, the rotor will accumulate dirt and grime from the air, which have some reflectivity." - I dunno. I think that you may be dismissing the problem of making a transparent rotor too casually. Sure, a moving rotor will accumulate some dirt and grime, but then again so does the fuselage of an F-22 or F-35 when they are flying.

Networked Warfare is another tactic that can be used to make helicopters "pseudo-stealthy". E.g. two helicopters hide behind different hills, one pops up its radar mast (which is small enough to be practically invisible) and relays target Information to the other which fires its missiles "blind" from behind the hill, so the enemy cannot observe where they are coming from.

@SamuelWeir Yes, both are somewhat reflective. The RAM on the skin itself doesn't absorb everything. The difference is that the rotating motion of the rotor makes it impossible to direct the reflection in a specific direction, the way fixed-wing skin can. You can make a stealthy object out of pure steel, no composites, no RAM. But it will instantly lose stealth if you start wiggling it around.

Well, there is the RAH-66. Stealth helicopters are difficult to make because of the radar cross-section of the blades. Even on the B-2 the fan blades are shielded so they don't create a massive return, and helicopter blades would create a much bigger return. Even with the Comanche's FENESTRON tail rotor, it's not very quiet. Helicopters are loud machines, no matter how you slice it. There may be quiet helicopters, but they're still pretty loud. To answer your question about obscurity, well, it''s like the B-2 or SR-71. Nobody wanted anyone else to know it existed because the tech was so valuable. Same with stealth helicopters.

Was not is in re Comanche. Cancelled about 16 years ago. (what a shame, but so it goes)


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Why does enabling engine sync on the MD-80 series trigger the "Master Caution" light?
I fly in a consumer PC flight simulator. Among my virtual fleet is a fairly faithfully-reproduced MD-83.
One fairly strange detail I noticed is that enabling engine sync (either on N1 or N2) in the overhead, while on the ramp, triggers not just the alert on the annunciator panel, but also the master caution light -- why is this? Engine sync isn't a particularly abnormal condition. Is this just an inaccuracy in the simulator?

Is this with gear down? The FCOM says "An ENG SYNC ON annunciation on the overhead panel will come on when landing gear handle is in the down position and the ENG SYNC selector is in N1 or N2 position."

That'd probably be it, thanks @Bianfable! I was toggling the switch on the ramp, and it looks like warning doesn't get toggled in the air with gear up.

I still don't know why the warning actually exists. The FCOM does not say. Presumably, it is bad to land with ENG SYNC on, but I'm not sure why, since it further says "The pilot can override the system at any time with throttles."

AFAIK The reason you aren't supposed to take off/land with prop or engine sync on is because if the leader engine fails, the follower engine will reduce power to match it, and that is really not ideal, especially low and slow, a la V1 cut or failure on short final. Try it out- fail the leader engine just after V1 and see what happens!

Nice answer, and makes intuitive sense the way you explain it. Please come back often - we could use more folks with Mad Dog experience here!


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Should I pursue aviation?
I am a senior in high school who has been interested/fascinated in aviation. I do not know whether to go to a 4 year school to fly commercial for UPS or Fedex or do charter planes or fly for companies like menards. I am not sure of the benefits of either one. I’ve heard some work more days and some work less days. I just want to be committed on this before I spend a boat load of money for school.

Welcome to aviation.SE! This is a good question but it isn't really a good fit for this particular site. We focus on questions that have a clear answer; open-ended and subjective questions work better on discussion forums, where you can get a range of opinions, ask follow-up questions and so on. As suggestions, you might try pilotsofamerica.com or reddit.com/r/flying. But do stick around and browse the training questions we have. The [tour] might be useful to help you with how the site works.

@user46189 I'd say spending the boat load of money is the measure of your commitment. If you hesitate with that one, this probably is not the job for you. Aviation is a highly competetive career, and it'll be years before you've paid for your training unless you were born loaded. There's no fast tracking for success.

You really need to talk to a few pilots.  People don't just get jobs with majors and a lot of folks don't like flying charter.  The job is tedious, only occasionally exciting, and ultimately very fickle. I found out that I really like flying but only on my schedule.  I don't like being evaluated every six months and I don't like flying when I don't want to.  Consider getting a high paying job and just buying a plane.


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Derivation of the Takeoff equation
I am trying to understand the derivation of some simple equations to calculate the neccecary takeoff speed for an airplane (under ideal conditions).

Which is taken from 15. Takeoff and Landing. My main issue is the derivation of the horizontal equation
To get from the first to the second line my guess is that we use
R = W - L   and    m = W / g

My main problem is understanding the details between what happens from line 2-3 and 3-4. I would love if someone could give me a brief explenation of the symbols used, what does S denote?

What don't you understand from 2-3? It's just rearranging the algebraic terms. #4 replaces lift and drag with the lift and drag coefficients.

We want to reduce (or simplify) equation 2, to get get dV/dt by itself, so we multiply both sides by g/W.  On the right hand side, this reduces to dV/dt.
On the left hand side it reduces to what is given in equation 3, but let me show you how:
 1. First term is T x g/W
 2. Second term is -D x g/W
 3. Third term is -u x g/W (W - L).  By the distributive property, we have -[u x g/W x W] + [u g/W x L]. The first part of this term simplifies to u x g.
Finally, I want to group like terms together.

The first term combined with the first part of the third term give me: T x g/W - ug = g(T/w - u).
The second term combined with the second part of the third term give me: -D x g/W + u g/W L = g/W(u L - D).

This yields equation 3: g(T/w - u) + g/W(u L - D) = dV/dt.
Finally, recall that L = 1/2 p S V2 Cl, with a similar equation for Drag, and we can substitute these directly into L and D to get equation 4.
S is the the wing planform area. Other sources use A, like this one: NASA - Lift equation.


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If pure N2 (not gas generator) is used for evacuation slides, why is CO2 still needed?
In the answers to Aircraft Emergency Evacuation slide inflation system: Why is a mixture of CO2 & N2 used instead of pure N2?, it's mentioned that the nitrogen comes from a gas generator.
Yet the bottles I'm familiar with are labeled as "55% of N2 and 45% of CO2", how so?
The maintenance manuals do not mention a pyrotechnic gas generator. So if pure N2 is used for evacuation slides, why is CO2 still needed?
Is there a technical reason, or simply a simplified labeling?

These a huge amount of detail on that other question, you should ask your questions as comments there rather than asking a duplicate question.

@GdD Someone just deleted my question in the comments because it was a question so...

You posted your question as an answer @electrotech, that's why it was deleted. In the future add a comment instead.

@GdD To be fair, you need a certain rep level to be able to comment.

Does this answer your question? Aircraft Emergency Evacuation slide inflation system: Why is a mixture of CO2 & N2 used instead of pure N2?

@selectstriker2 Not really because he is saying that the N2 comes from the pyrotechnic gaz generator but the latter is not mentioned in the maintenance manual... It is clearly written that the bottle needs to be filled with half CO2 and half N2. So I don't really understand why we need this amount of gases

Peter's answer literally says "Now we can answer your question: Nitrogen is contributed both by the pyrotechnic gas generator and by the compressed gas and is produced from the dissociation of nitrous oxide (N2O)." Plus it answers every one of your questions pretty exactly.

@selectstriker2 That doesn't answer my question. If the N2 is produced by the dissociation of N2O, why the bottle is also filled with 55% of N2 ? Because as it is written, the dissociation produce N2 so we don't need to add more N2

@selectstriker2 that is the question he explicitly stated that he'd read and still didn't understand.

Nevermind, i'll go away without my answer, I just seem to bother you

@electrotech what maintenance manual are you referencing?

I still think this is a duplicate. From the accepted answer to the other question: "Carbon dioxide is added to dampen the exothermic dissociation and absorb some of the heat produced.".


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Do the white lights on the chimneys of the Tokyo area incineration plants serve an aviation purpose?
At least 新江東清掃工場 (Shinkoto Incineration Plant), 江戸川清掃工場 (Edogawa Incineration Plant) and 市川市 クリーンセンタ (Ichikawa Clean Center) ー blink with white lights in series.
I know that the red lights of other buildings and structures are for air traffic safety.
But these plants are relatively low. Their chimneys have the lights attached. And as said they blink white.
Yet the near port probably has no use for these lights.
I suspect this is for air traffic. Is this some sort of guidance system marking flight vectors or something?

I'm guessing that very few of us read Japanese, could you translate that part into English? Is it the name of some plants? Usually they have English equivalents.

it's HND (Haneda Tokyo)

Since you are writing your question in English, on an English language website, what is the purpose of embedding Japanese language characters?  Does this contribute anything to our understanding of the information you seek?

These are place names, you'll need them to find them in Google maps.


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In the Cl vs Cd graph, Why the drag coefficient decreases initially with the small increment in lift coefficient?
Cd = Cdo + KCl2
Total drag = Parasite drag + Induced drag
When  Cl = 0 , Cd = Cdo
Increasing in Cl introduces another component of drag (Induced Drag) and therefore the total drag Cd should be increased. But, Why in this plot Cd decreases initially with the small increment in Cl?

I don't think the origin of the graph is at (0,0), so your question may be based on an incorrect premise.

Can you please add the source of the image?

I'm not convinced that the origin is not at  0,0.

Possible duplicate of Is drag coefficient lowest at zero angle of attack?

Think about it this way.  Assume a cambered airfoil, not a symmetrical airfoil.  Assume the wing has zero twist and zero incidence.  To place the wing at the zero-lift angle-of-attack (zero CL), the fuselage and wing will have to fly at a somewhat nose-down pitch attitude relative to the airflow.  This does not yield the lowest possible drag coefficient, regardless of whether we are looking at the wing, the fuselage, or the whole aircraft.
Bear in mind that in this part of the flight envelope (near the zero-lift angle-of-attack), drag is dominated by profile drag, not induced drag.
Related:  Is drag coefficient lowest at zero angle of attack?
Now imagine using the same data to generate a graph of CD versus airspeed, assuming Lift=Weight.  What would that look like?  Also what would a graph of Drag versus airspeed look like?

https://aviation.stackexchange.com/questions/62357/is-drag-coefficient-lowest-at-zero-angle-of-attack

Right, now incorporated, thanks.  I guess it's a duplicate question or nearly so.  Zero a-o-a and zero-lift a-o-a are not exactly the same thing.

The well-known 3D drag model of:
$$C_D=C_{D_0}+KC_L^2$$
only holds well for high aspect ratio, flap retracted wing configurations (source: ESDU Item 74035).
With flaps extended, a better model would be (source: ESDU Item 97002):
$$C_D=C_{D_0}+AC_L+BC_L^2$$
If you rearrange the terms a little bit, you get the model shown in your graph:
$$C_D=C_{D_{min}}+K(C_L-C_{L_{mindrag}})^2$$

I don't think the flaps are the cause of the $C_{D_{mindrag}}$ and $C_{D_0}$ discrepancy. This phenomenon also occurs when the flaps are retracted (due to the reasons mentioned by quiet flyer). Also, it's not clear to me how the rearrangement of terms (in which equation, the one with the flaps?) should lead to the final result

Source: INTRODUCTION TO FLIGHT - John D Anderson

might be copyright issues just posting part of a book like that


text:
Does deploying flap change the aspect ratio of the wing?
Specifically when a Plain flap is defleted the geometry of the wing will be altered (Camber is increased). Does this have any effect on aspect ratio? What's the connection between Camber and aspect ratio?

There is no connection. Aspect ratio is based on a projection of the wing in the x-y-plane. Camber is in z direction.

Short answer: it depends. Some flaps extend backward (increasing chord and thus reducing aspect ratio), other don't. See The section about type of flaps on the Wikipedia page so that you can evaluate if you want to restrict your question to specific type of flap.

The aspect ratio of a wing is defined as:
$$ \text{AR} \equiv \frac{s^2}{A} , $$
where $s$ is the wingspan and $A$ is the total wing area.
Some types of flaps are not only increasing wing camber, but also the wing area:

The Fowler, Fairey-Youngman and Gouge types of flap increase the wing area in addition to changing the camber.

(Wikipedia)
Since the wingspan is unchanged, this will decrease the aspect ratio. Other types of flaps, like a plain flap, are not changing the aspect ratio.
You can see the flap types that change the aspect ratio in the following image (modified from Wikimedia):


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What is a ram recovery point?

As aircraft speed increases, thrust tends to decrease somewhat; as the aircraft speed reaches a certain point, ram recovery compensates for the losses caused by the increases in speed. The inlet must be able to recover as much of the total pressure of the free airstream as possible. As air molecules are trapped and begin to be compressed in the inlet, much of the pressure loss is recovered. This added pressure at the inlet of the engine increases the pressure and airflow to the engine. This is known as “ram recovery” or “total pressure recovery.” The inlet duct must uniformly deliver air to the compressor inlet with as little turbulence and pressure variation as possible. The engine inlet duct must also hold the drag effect on the aircraft to a minimum.

Source: https://www.flight-mechanic.com/turbine-engine-inlet-systems/
But how the thrust decreases, and how Ram pressure increases?

FYI, on the same site, this other page provides additional details (last third), e.g. this graph.

The thrust ($T$) of an engine is $T = Q(V_{out}-V_{in})$ where $Q$ is the mass flow through the engine, $V_{out}$ is the exit velocity of the exhaust gas and $V_{in}$ is the speed at which the air flow enters the engine inlet.
The mass flow $Q$ is a product of volume of the air sucked by the engine and air density.
When engine is producing considerable thrust, the exhaust speed ($V_{out}$) is the speed of sound. So for the sake of simplicity, we can consider that more or less a constant value. When aircraft is stationary, like at the beginning of the takeoff roll, $V_{in}$ is zero. As the aircraft begins to accelerate, $V_{in}$ increases.
Air is compressible gas, but at low speeds compressibility is negligible. As the forward speed increases, the $Q$ is constant for a given engine RPM, the $V_{out}$ is constant and $V_{in}$ increases, the net result is that engine begins to lose thrust.
But as the speed increases past about Mach 0.2, which equals approximately 130kts at sea level at standard day conditions, the air being packed (rammed) in the engine inlet begins to compress due to increasing pressure. That means that  the density of air in the inlet increases. As density increases, and if we assume constant RPM (and thus volume) the mass of the air flowing through the engine increases.
At some point the increased mass flow $Q$ has compensated the loss of velocity change in the engine and thrust has reached its stationary engine value. This is the ram recovery point.

In class, we haven't talked about it as a specific recovery point.
But as a gradual increase in thrust based on airspeed.
Q: How  does the thrust decrease?
Airspeed thrust decreases as the altitude increases. (High altitude = low pressure/less air)
less air = less thrust
Q: How does ram pressure increase?
Ram pressure increases when you fly into the air molecules fast. This happens because you push the air molecules together.
Q: What is the ram recovery point?
I had a chat about this with my turbine engine teacher, which changed my understanding on the subject.
 Though you could see the ram recovery point as the moment the airspeed effect on thrust and and the ram effect on thrust are equal.
Here is a graph showing the result of ram recovery speed, and the point where the forces are equal.
(From the Dale Crane part 66 Powerplant book).
You reach the ram recovery point as soon as the inlet pressure (ram pressure) is greater than the ambient pressure.
In an idle engine run on the ground, the engine has to suck in air. As soon as the engine doesn't need to suck, and air is forced into the inlet - you have reached the ram recovery point.


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What exactly is engine design speed?
Any further information is welcome, this is what I have learned so far:
Engines are designed to operate at a certain RPM (N1). The compressor consists of small 'wings' which are fixed in position, so the angle of attack will vary only by engine speed and forward speed. The engine design speed should be 85-95% N1 for most turbofan engines, this results in an optimum airflow through the compressor. Above or below a higher risk exists for an engine to stall.

The "wings" inside the engine are called "vanes". They are not as "fixed" as you may think, they float in the hub to allow for thermal expansion/contraction, but they do not vary the angle.

correct thanks, term fixed was referring to the angle of attack.

@ymb1 I don't think a lot of turbofan engines used in commercial aviation have variable stator designs, but yes, they do exist (mostly in a military context to avoid compressor stall/surge). GE has a 10-stage with the first few being variable but I'm not sure it's used anywhere but military aircraft, or if its still in the development stage.

Thanks, didn't know they made it into production engines.

@ymb1: I was under the impression that essentially all modern turbofans, even civil ones, had VSVs.

@ymb1: Actually, I was responding to your comment that only "many" civilian turbofans had them.  :-)

Engine design speed is the shaft RPM and vane pitch which yields the rated thrust that the engine was originally intended to furnish. This will typically be quoted for takeoff at or near sea level at a standard ambient temperature (including a time limit to control maximum temperatures inside the engine) and for optimum cruise conditions at altitude (which will include specification of nominal fuel burn rate in pounds per hour).


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Do high wings aircraft create an additional vortex in the middle?

At the same time, the air on the upper surface has a tendency to flow in toward the fuselage and off the trailing edge. This air current forms a similar vortex at the inboard portion of the trailing edge of the airfoil, but because the fuselage limits the inward flow, the vortex is insignificant.

_(FAA Pilot’s Handbook of Aeronautical Knowledge, chapter:5 ,Pg 8)
This got me thinking if there is no fuselage, as in the case of high wing aircraft, the inward flow is not restricted anymore. So what will happen to them, do they form new significant vortex? Or there will be another pressure drag?
P.S. I could not find any image of a high wing aircraft with vortices in middle. So if you get one (if it happens so) please link it in the answer. I am too eager to see.

Short answer
If there are two wingtips, there are two vortices.

Vortices - summary
There are only two vortices, produced due to the very existence of ends to the wing - the wingtips. Pressure distribution that exists over the wing ceases to exist beyond the wingtips (since there's no wing beyond the wingtips). To understand this in detail, first consider 2D flow.
In 2D, the low pressure region on the upper surface of the wing sucks the air from all four directions: It pulls the air from ahead, causing it to accelerate and upwash. It pulls the air from behind, causing it to decelerate (this is the pressure recovery). It pulls the air from up above, causing it to downwash. And in an attempt to pull the air from underneath, it pulls the wing up, thus producing lift.
In 3D, this low pressure also tries to pull the air from sideways. On an infinite wing, the pressure distribution is homogeneous spanwise, and so there is no spanwise flow. On a finite wing, there are two wingtips, beyond both of which there isn't a similar pressure distribution. This causes the low pressure region above the wing to pull in the air from beyond the wingtips. The high pressure region underneath the wing similarly expells air outboard, beyond the wingtips. This is what causes wingtip vortices.

High Wing Configuration
On a high wing aircraft, on the part of wing above the fuselage, there's still a region of low pressure (and even if that portion on its own didn't have a low pressure region, the pressure from the consecutive parts of the wing will be carried over there). More importantly, this part of the wing is not like a wingtip, and shouldn't be treated like one; there is only one wing, and two wingtips (the "left wing" doesn't terminate at the fuselage, it extends from there all the way to the right wingtip). Since there are no wingtips in the middle, there are no vortices.
You can think of it this way: both the "left wing" and the "right wing" try to form a vortex in the middle; the two POTENTIAL vortices are equal in magnitude but opposite in direction. The two potential vortices completely annihilate each other, and so there's no vortex in the middle. (you can apply this idea at any point on the span, except that the two opposing vortices won't always be equal in magnitude; they are only equal at the middle of the span. If you do this for every point, the end result will be the two wake vortices as we know them)
As far as your handbook's claim goes that the fuselage acts as an end plate, restricting the vortex formation - yes, that's technically true: the fuselage will act as an end plate for the left wing¹, largely inhibiting it's vortex. But the right wing will do a better job at this, eliminating the left wing's vortex altogether, as described earlier. If anything, fuselage is actually making the situation worse by not letting the right wing properly do it's job (that's why the two small vortices form in first place).
¹same is applicable for right wing

Thank you for such an explanation. I did understand the concept of vortices. But the point where two equal or almost equal in magnitude inward flow carrying some momentum meets and cancel each other without any energy transformation like drag or something is a little hard to imagine, may you please enlighten about this too. Thank you.

@NoorulQuamar If a wing was to be cut in half, two entirely seperate wings will be obtained - both of which will independently produce two tip vortices (four in total). As we attempt to join the wings back together, we will observe that the two vortices at the centre will cancel out (because they are counter-rotating).

Actually since the half-wings are joined, the pressure discontinuity which existed at the edges of both wings no longer exists. Earlier the low pressure above the wing ended at the edge, and was able to pull air from beyond the edge; but now, the low pressure extends beyond the edge, since there's the other wing beyond the edge. This means that there is no pressure difference across the edge and so the vortex cannot form

There is no energy transformation (drag) because there is no pressure difference. If there was a pressure difference, like there was when the wings were not connected, then the vortices would form and then there will be drag. Now there is pressure equilibrium at the centre, therefore that particular drag is avoided.

FAA handbooks are known to make scientifically inaccurate statements on occasion. This may be one of those times. What they probably intended is two separate wings magically flying with a gap where the fuselage would normally be, rather than a continuous span like a B2 flying wing. This 'two wings with a gap' situation would indeed have inboard vortices, because it would have inboard wingtips. However, including such a contrivance in a practical manual may be more ridiculous than if they had made a simple scientific error.
Anyway, flying wings ( Even two in close formation ;p ) produce vortices at the wingtips as the air is accelerated by the difference in pressure between the lower and upper surfaces. Air has mass [inertial resistance] and it is very compressible and extensible; so it creates a pressure gradient that has effects far inboard from the wingtips. The amount of effect (primarily spanwise flow) is a continuous [non-linear] gradient from maximum near the tips to zero at the center.
There are cases where a sudden sectional change can create smaller additional vortices, such as the outboard end of a large extended flap. These are caused by the same type of pressure effects as those at the wingtips.

Where does air flow?

Air flows from high to low pressure

Underneath the wing, air pressure tends to be higher than ambient.
The fuselage acts as an endplate, and pressure tends to leak out the wingtip and aft.  The front is a ram.
So, this "engine on a half-shell", or sail, has its highest pressure next to the fuselage.  It gets better with flaps.

"because the fuselage limits the inward flow, the vortex is insignificant"

Essentially, this is saying the fuselage acts as a "fence", or endplate.  This idea has been carried forward with little fuselages on wingtip, know as "tip tanks".
The thought seems to be an extension of the logic that each half wing ends in 2 places.  In reality, the fuselage has lifting properties as well.


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How are Fan pressure of and Static pressure in the turbofan engine related?

If the fan air is ducted to the rear, the total fan pressure must be higher than the static gas pressure in the primary engine's exhaust, or air will not flow. By the same token, the static fan discharge pressure must be less than the total pressure in the primary engine's exhaust, or the turbine will not be able to extract the energy required to drive the compressor and fan.

Source: Aircraft Gas Turbine Technology by IRWINE TREAGER, Chapter 2, Section: Turbofan Engine
I am not able to understand:

Why total pressure of fan is being compared with static pressure of engine?
Is bypassed air which is ducted to rear is involved in this operation? And how the pressure difference is necessary to prevent backflow when both sections do not meet at all?
How turbine energy extraction depends on static fan discharge pressure? And is the discharge pressure involves ducted air pressure?


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Why Lift-induced Thrust Required(Tr) decreases and Zero-Lift Tr increases with the increase of velocity?

As the minimum Tr is at the point where L/D (lift/drag) ratio is the most. And Lift-induced Tr depends on the Coefficient of Lift, (Cl)^2 and Dynamic Pressure (q) which also increases as velocity increases. But why it is decreasing?
And how Zero-Lift Tr depends on velocity and is affected that much when its coefficient is a constant?

I am offering https://aviation.stackexchange.com/questions/36062/ as duplicate. It asks generally why the induced drag formula looks like it does, and contains a detailed derivation.

There are also How is the induced drag calculated for a wing with elliptical planform? and Is the induced drag independent of wing span? with further derivations of induced drag.

@JanHudec, thank you but I had this formula already. I am grateful that you reminded me it was a level, unaccelerated flight. What I needed is a formula for zero-lift Tr --while its coefficient is a constant, its formula is not: it actually has a cube of velocity. as said by @zaitcev

Let us continue this discussion in chat.

This happens because Angle of Attack required to maintain flight at a given weight decreases with the increase of speed.
As for the zero-lift Tr, also known as "parasitic drag", while its coefficient is a constant, its formula is not: it actually has a cube of velocity.

Yes, I am aware of the fact that AOA decreases, but how it relates to increasing Zero-lift Tr? I am sorry if it is very simple yet I am not understanding.
Edit: Found your reason later, may you please direct me to the formula. Thanks.

re "it actually has a cube of velocity."-- can you expand on this?

There is no cube of velocity in any drag. The aerodynamic formulae (except when dealing with compressibility) only include velocity in dynamic pressure terms $q = \frac12\rho v^2$, so velocity can only appear in them in even powers.

Thanks to @JanHudec for reminding me that the graph is for a level flight.
To keep the flight level at low speeds, the Angle of Attack(AoA) needs to be increased, meaning Cl (coefficient of lift) must be increased and so the lift-induced drag ( proportional to Cl^2) is high at low speeds.
With the increase of speed, AoA decreases to keep the flight level, hence, Cl and so lift-induced drag and its Thrust Required (Tr) decreases.
Coming to Zero-lift Drag or the parasitic drag. At low speeds it is low. But at higher speeds, there is an introduction of wave drags,
"parasitic drag increases because the fluid is striking the object with greater force, and is moving across the object's surfaces at higher speed. As speed continues to increase into the transonic and supersonic regimes, wave drag grows in importance."-Wikipedia.
Therefore, in the case of Zero-lift, the increase of velocity is not much of the reason (by this I mean, not the proportionality with the square of velocity), but the increase of wave drag is. And so there is not much effect on Induced drag with the increase of velocity


text:
Why are there so few carrier capable delta wing aircraft?
When looking through various carrier aircraft, I noticed that there seemed to be very few with delta wings. While initially chalking this up to higher stall speeds necessitating higher approach speeds on supersonic deltas, I nevertheless noticed that this seemingly doesn't apply to the Dassault Rafale, nor the HAL Tejas.

Is this due to configuration? Would the addition of canards or conversion to a tailed delta permit carrier operations? Or is this not aerodynamic but more based off cost?

I'm assuming the first sentence is intended to mean ALL delta wings regardless of whether there is also a tail or canard?  Or do you only mean tail-less canardless deltas and want to know why that configuration is so rare?  If so the question could benefit from a clarification.  The list of carrier aircraft with delta wings AND tails (or canard wings) would be rather long.

Are you asking for a country in particular, or any country?

@quietflyer I meant delta winged aircraft in general including canard or tailed configurations, I wasn't aware that there were that many with delta wings, seemingly jsut the F4D, A-4, Tejas and Rafale, the rest all have straight, swept or trapezoid wings.

@RonBeyer In general really, although the planned carrier capable Mirage III specifically was the one that got me interested in the question

I really appreciate seeing question about aircraft manufactured in non-western country (the HAL Tejas). It is really rare, especially on aviation.SE.

@quietflyer I've never thought of the Phantom as having a delta wing (although I may well be incorrect there). I've always understood a delta wing as being triangular like the letter delta, relying on Wikipedia for those that fall into the overlap between cropped delta and swept. Based off that I'd say any which Wikipedia calls a delta, I'd count as a delta for this question.

@G01 -- ok, your observation may be more correct than I thought, A-4 and F4D immediately popped into mind but perhaps you are right that it is not so common. I assumed Super Eternard had a delta wing too but I was mistaken, and likewise with a few others. May depend in part on how much modification disqualifies it from being a delta-- e.g. F-4 Phantom?  (Reposted to fix typo)

Delta falls out of favor in general for the current generation of fighters. Proportionally there are as many deltas on carriers as on land.

The reason may be change in tactics.  The 4FD was designed to intercept aircraft, and as a point defender and dogfighter, it would have been very successful in this role.
But the WW2 scenario of planes appearing overhead to drop bombs changed to planes launching cruise missiles from farther away.
Carrier aircraft need to go further out to intercept them, carrying as much ordinance and fuel as they can.  Deltas simply do not do this as well as even lower aspect straight wings, because the design does not lift as efficiently for the amount of drag produced.
Lack of stealth would add to their modern-day woes.
The F4D, in its day, possessed many of the advantages the WW 1 triplane did in maneuverability,  this time with cannon and missiles.  Like the triplane, it became obsolete.

Maneuverability around a carrier is still in demand today. Remember that the Boeing X-32 started as a pure delta and grew a horizontal tail later when Navy demands could not be met. And this delta was stealthy, so that is not the reason.

The F-35C does have more wing.

Just speaking for the case of Naval variant of Tejas, it has the "Leading Edge Vortex Controller - LEVCons' as seen from the photograph, which help lowers the landing speed and maintain the landing distance around ~ 200 meters as per the following article :
https://swarajyamag.com/insta/lca-tejas-naval-prototype-with-new-levcon-flaps-arrestor-hook-mods-to-begin-final-shore-based-trials


text:
How are Aircraft Noses Designed?
Look at the Learjet, Embraer and Boeing noses which are kinda pointy - drag efficiency aside, how do they come to the final design of the nose?
I’m talking from an aesthetic point of view - how do they design them to be aestheticly pleasing as they are?
By hand, CAD software or other? I don’t think it’s random guesswork I produce what looks good and what doesn’t - do my question is: how do they make them noses?
I don’t think in the early days of the 737 they had advance mesh modeling like we do now!

They didn't have advanced mesh modelling, but they did have a physics book, wind tunnels and a lot of experience. Experience was very valuable, perhaps even more than it is now.

Of course 'mesh modeling', no matter how advanced, would not come into play here. Parametrical modeling would be used where the surfaces are described by mathematical equations. Meshes are only created as an approximation of the actual surface, and used for simplicity's sake in areas such as CFD, RCS prediction, and of course visualization.

A fascinating thing is that, I believe for new ideas on supersonic aircraft (to counter the boomy thing), a really REALLY long point is called for.  Cool !

@Fattie supersonic flight is a very different regime with very different conditions as opposed to typical commercial cruise at ~M0.85. The pointed nose serves a purpose beyond the "usual" aerodynamic functions of a rounded commercial nose cone, owing to managing shockwaves in supersonic flight.

related: https://aviation.stackexchange.com/questions/12915/why-do-boeing-and-airbus-have-distinctively-different-nose-designs

Also related: https://aviation.stackexchange.com/q/30112/1677

@verandaguy , yes, that's correct. They do seem to be literally building this BTW:  https://www.newscientist.com/article/2165683-nasa-is-trying-to-build-a-supersonic-aircraft-without-the-boom/

Aesthetics are by far and away the last thing aircraft designers are optimising for. A better question might be, why do aerodynamically-efficient designs look pleasing, but that would likely be an art or psychology question and off-topic here.
Since aircraft could travel at appreciable speeds, designs have focused primarily on drag reduction. Smooth curves and teardrop shapes on the leading parts of shapes help with this. Before 3D fluid modelling software, this kind of design would be achieved using a combination of fluid dynamics computations done by hand, wind tunnel modelling, and likely a bit of intuition.

If aircraft designers aren't optimizing for aesthetics, why do Airbus noses look different from Boeing noses? (more rounded vs more pointy) - shouldn't they have arrived at the same "most aerodynamic" shape?

"why do aerodynamically-efficient designs look pleasing?"
Probably because we associate sharp looking objects as threatening and painful to touch, and smooth looking objects as inviting and pleasant to touch.

@Skyler you'll notice that in newer ground-up designs (like the 787 and A350), the design elements tend to converge: smoother noses overall, highly raked wing-tip devices, etc.. Historical differences stemmed from the age of parent designs (737s are ~50 years old as a family vs A320 which is about 30) which were informed by both available materials (steel/alum vs composite/carbon etc.), simulation methods (wooden models vs CFD), and available production methods and logistics (local generalized vs global specialized assembly).

I'll also refer to @JohnK's comment about component integration including windshields to back up the point about available materials and production processes.

Something that is true of both aircraft and sailcraft: they are designed almost complete without regard to esthetics and end up being the most beautiful items ever manufactured.

@Malvolio I think that is because their form is constrained. Very few of our physical artefacts are as rigorously formal like, say, a poem is or some pieces of music.

It's mostly because of one of the most attractive things about aviation (shared with the marine world to a large degree).  Form and function, aerodynamic necessity and aesthetics tend to coincide.  Aerodynamically efficient shapes also tend to be the most artistically pleasing shapes.  Airplanes that fly well tend to look good.
Beyond that, you have the need to fit the radar antenna of whatever weather radar system you want to use, while achieving an efficient curvature and fineness ratio appropriate to the airplane's speed range.
Then you have to integrate the windshields into this. Flat windshields are way cheaper than curved ones, but are more draggy.  And, going back to my original point, the curved ones just look way better. But 50 years ago a multi-laminate curved windshield that could handle 8 psi pressure differential would cost a fortune, so all the early airliners had flat windscreens and side panels.
It's a lot cheaper to make complicate curved windshields today so you see them pretty much on all new designs.

When painted properly, even a Skyvan can look nice, no matter the nose.

Good point, but I don't think the Skyvan is an ugly airplane in the first place.  More like utilitarian, but still with various pleasing forms in key places.

This is exactly correct.  Race cars and aircraft look good because they are purposeful.

A quote from my father, who flew with the RAF back in the 1950s and 1960s.  "If it looks right, it flies right."  A few examples of this are the Supermarine Spitfire, the De Havilland Mosquito, the Hawker Hunter and the English Electric Lightning.  All of which are really nice planes to look at, and very well behaved in the air.  This does NOT however mean that not attractive is bad: I will cite the A10 Thunderbolt as a counter example.  So ugly it's called the Warthog, but pilots that fly it really are in love with it - it's just a great plane to fly.

@dgnuff: if a skyvan qualifies as "not ugly", then an A10 is nothing short of gorgeous.

@JerryCoffin A-10s are horrible looking... from the front... when their gun is spun up... and you don't have a few inches of steel armor to hide behind.  Then they look hideous, at least for a few brief moments.

The shape will be driven almost purely by functional concerns.
How are those functional concerns tkane into account? It's not aircraft noses specifically, but the paper "Geometric Programming for Aircraft Design Optimization" explains how geometric programming, a form of convex optimization, can be used to efficiently extract optimal shapes and trade-off curves for aircraft wings.
CVXPY, an open-source optimization framework in Python, now has the capability to manipulate geometric programs. This makes it easy to experiment with such designs yourself.

Can you include the main points of the paper here instead of just linking to it?


text:
Why do all the rings of MH370's transmission locations have the same origin point?
This question originates from the video "The Vanishing of Flight 370", by Lemmino.

At 6:00 he says

While these transmissions did not contain any information about the flight's position, investigators were able to measure the distance between the satellite and the aircraft at the time of each transmission, based on how long it took those transmissions to be sent and received. This generated seven rings of possible locations, from which seven of these transmissions are thought to have originated.

And this diagram is displayed:

But something about the visual strikes me as odd: don't satellites move? I thought they had to keep orbiting or they would crash. So why do these rings of transmission locations, taken over a roughly six-hour period, all show the satellite above the exact same place? Clearly something about my reasoning must be wrong if the investigators never considered it.
(I'll understand if this belongs more on Space Exploration stack exchange.)

That looks more like a visual aid than something that's supposed to be looked at critically. But many satellites are kept in geosynchronous orbit so they can stay roughly over the same spot by orbiting at the same speed of the Earth's rotation.

Maybe geosynchronous, maybe geostationary. Either way, I'd agree that the for-TV graphic may well contain over-simplifications. The guys went pondered on this for weeks & weeks undoubtedly considered the satellite's position & movement relative to the earth, even if the video's graphics didn't.

There were two incoming SAT phone calls at 18:39 UTC and another much later in the flight at 23.13 UTC. All incoming calls were initially acknowledged and then abandoned by the SDU for voltage irregularity, or signal stability issues. its also evident that whilst we know power returned to MH370’s SATCOM system at 18:25 UTC, the ACARS system itself never resumed operation. What we’re left with is pings sent to the aircraft asking if it’s still online and ready to receive

Because the Inmarsat 3 series satellites that were used to transfer the ACARS system data are geostationary. They orbit at a speed and altitude that causes them to stay in the same location with reference to the earth. The signals were received specifically by the Inmarsat 3 F1 satellite which was, at the time, geostationary over the Indian Ocean.

Notably, "Inmarsat owns and operates 13 satellites in geostationary orbit 35,786km (22,236 miles) above the Earth," per https://www.inmarsat.com/about-us/our-technology/our-satellites/.


text:
Why is MH370's location data from the Malaysian military considered reliable if the altitude data isn't?
This is another question originating from the video "The Vanishing of Flight 370", by Lemmino.

At 13:52 he says

The data recorded by the Malaysian military does indeed contain altitude information, but it is highly inconsistent. In fact, a Boeing 777 is incapable of performing the extreme altitude fluctuations recorded. At one point, the aircraft exceeded its maximum operating altitude by more than 15,000 feet, before making a 50,000 foot nosedive in less than a minute. Attempts to recreate these maneuvers on a flight simulator were unsuccessful, and thus the data was deemed inaccurate and unreliable.

If the radar's data about the flight's altitude is untrustworthy, why is absolute trust placed in the radar's data about the aircraft's position? What force could cause the altitude data to be corrupted but not the location data?

That's just how radars work.
A radar works by sending pulses of radio waves and measuring how long it takes for them to be reflected back. If you know the speed of light within air, you can then calculate how far the radio waves have traveled, and from that, you know that the aircraft is half that distance away from the radar.
This is not terribly useful, though. You could either build an omni-directional antenna, which sends and receives pulses in all directions. Then you can tell the distance to every aircraft around you, but you can't tell where they are. Or, you could build a directional antenna, then you know that any reflection you get must be coming from the direction the antenna is pointing … but you can only detect aircraft which are in the direction of the antenna.
To solve this problem, radar combines the two approaches above: it uses a directional antenna which can only detect aircraft in one direction, but this antenna is put on a rotating mount. This way, you can tell both the direction and the distance of an aircraft.
One further trick you can do is measure the Doppler shift of the return signal. This tells you the aircraft's relative speed component in the direction of the radar. Note: it does not tell you the aircraft's actual speed, nor does it tell you its velocity vector. You can only tell how fast the aircraft is moving towards or away from the radar antenna – any movement sideways will not be detected.
As a last trick, a computer can integrate measurements from several rotations of the antenna. This allows you to tell which direction the aircraft is moving and how fast.
You will note that one thing was not mentioned here: altitude. A radar can measure distance from the radar. If it is rotating, it can tell direction relative to the axis of rotation. If it measures Doppler shift, it can measure speed along the radial. If it integrates measurements over a period of time, it can measure velocity relative to the axis of rotation.
However, all these measurements are always relative to the axis of rotation, and usually, radars rotate around a vertical axis, thus they "scan" horizontally.
In other words: a radar cannot measure altitude. So, it should not be surprising that altitude data from a radar is unreliable.
Note: all that I wrote here applies to so-called primary radar. Modern air traffic control relies on so-called secondary surveillance radar (SSR). This is really a misnomer. SSR doesn't work like a primary radar at all, it is actually a simple wireless data-link protocol. The "radar" sends out an interrogation message and the aircraft replies with a data message. Which data is included in this message differs based on exactly what protocol the radar and the aircraft's transponder support.
The response includes at least some form of identification (either the 16 bit transponder code or the 24 bit ICAO address). In almost all cases, it also includes the aircraft's altitude. It may also include position, direction, and velocity.
However, crucially, it requires the aircraft to be cooperative. If the pilot turns off the transponder, SSR will not work.

SSR isn't a complete misnomer, it still has some radar aspects in that the round trip time between sending out an interrogation and receiving a transponder reply gives the range and the direction of the ground radar antenna gives the bearing. The antenna still scans rotationally (it's even often mounted on top of the primary radar) and this is crucial for SSR to be able to match returns to targets (where a simple data-link could conceivably use an omnidirectional antenna instead).


text:
Can a normally poorly handling airframe with high moments of inertia actually have superior flight dynamics under autonomous control?
I’m not a pilot but know a thing or two about flight dynamics. As you know there are currently quite a few aircraft under development designed for low-and-slow and (eventually) autonomous flight. These are often referred to as air taxis, eVTOL planes et cetera. Their engineering may require a lot of new thinking but perhaps not. This relies in part on the answer to the question whether they’re designed to be optionally autonomous or exclusively so. My question pertains to aircraft that are NOT to be controlled by humans or only in the case of an emergency, but even then only indirectly through an advanced fly-by-wire system, one probably more advanced than those in modern fighter jets which would be impossible to fly without FBW.
The autonomous air taxis will be more affected by atmospheric turbulence than commercial high-and-fast jetliners and turboprops, and they’ll have to dynamically compensate for this to ensure a smooth and safe ride. To achieve this requires continuous adaptation relying mostly on gyroscope measurements coupled to small to moderate steering actions under autonomous control. Responsiveness is key here and that’s what computers are good at.
So let’s say we’re designing an airplane that’s to be autonomously controlled at all time (obviously with sufficient redundancy). Might for example a plane with a high moment of inertia provide a smoother and at least as safe ride as a “properly” designed airframe designed for manual control? Or do you think the old rules still apply: low(-ish) rotational inertia provides superior handling and passenger experience.
My thinking is that a flight computer that’s programmed to stay within rationally determined limits (to prevent spins et cetera) can fly a plane not just safely but smoother as well. This is akin to how very large ocean vessels are hardly affected by waves and winds compared to small ships. The small vessel is likely to have superior handling characteristics under optimal conditions and yet you'd be much rather off sailing on the large one even if both ships are able to stay afloat in extreme weather.

Questions asking "do you think..." are explicitly off topic as opinion based.

The B2 is an example of a touchy-to-control design (flying wing), being made stable by computer system control. It's not for passenger comfort, of course, but for maintaining flight stability while remaining in specific flight corridors, which an be very small, when on mission.

The underlying assumption of your question seems to be that a higher control bandwidth always leads to better disturbance suppression. Since a higher control bandwidth is easier to achieve with low-inertia systems, aircraft with low inertia would have better disturbance suppression.
This assumption can work in theory, especially for SISO (single-input, single-output) systems. However, for aircraft, we are essentially dealing with MIMO systems. While we can achieve near-perfect disturbance suppression for one degree of freedom (e.g., zero variation in altitude), this will inevitably lead to large excursions of other degrees of freedom. For example, disturbance suppression in altitude will lead to large variations in pitch as we try and keep AoA constant with every gust. So, for a MIMO system like an aircraft, we have no choice but to formulate a cost function of what we consider a 'smooth' ride, and balance the control scheme across various degrees of freedom.
Since we are considering a balancing act, it then makes sense to tune the internal dynamics of the aircraft such that a smooth ride is naturally achieved. So that means: high inertia for degrees of freedom experienced by the passenger, and low inertia for degrees of freedom not felt by the passenger. In practice: larger fuselage inertia and smaller control surface inertia.
Unsurprisingly, this is how it works in practice: larger aircraft are generally more comfortable, whereas at the same time, degrees of freedom not felt by the passengers are exploited in modern aircraft for gust suppression. For example, wings are more and more dynamically decoupled from the fuselage by highly flexible composite wings. Ailerons are split in smaller sections for gust suppression, you may find videos of "dance of the ailerons" for modern aircraft.

I'm not sure if this answers the question because IMO the question isn't clear, but it's still informative.


text:
What happen if the glide slope beacon removed for maintenance?
it about aircraft navigation system (ILS), when the aircraft about to landing but unfortunately the glideslope beacon being removed.

The aircraft is about to land but unfortunately the glideslope beacon is removed Just kidding!

Instrument Landing Systems routinely undergo maintenance and it is common to have the Glide Path unserviceable while the Localizer is still serviceable.
An aircraft can still make an approach and land, but the approach is now classified as a Localizer approach. The weather limit minimums are increased from the usual 200’ above ground, to approximately 300-500’ above ground. (depends on local geography)

The question is a little ambiguous. This is assuming you can get clearance for the approach and that it has not been NOTAMed as completely out of service.
If you are asking about the procedure if the glideslope vertical guidance transmitter is inoperative, in most cases you treat the approach as a LOC localizer or LNAV only approach. You would use the LOC/LNAV only procedures, criteria, step-downs, and minimums. You would have an MDA(H) (Minimum Descent Altitude or Height) instead of a DA(H) (Decision Altitude or Height).
If you are asking about the procedure if the marker beacon(s) is/are inoperative, you would use one of the alternate means of identifying the Final Approach Fix. This could be done utilizing DME or GPS. In some cases, it is marked on the Instrument Approach Procedure chart as an intersection of a VOR Radial. Otherwise, you would follow the glideslope vertical guidance all the way down to the DA(H). Make note that the DA(H)/MDA(H) may be adjusted if you can not identify certain points, or if the altitude can not be verified due to unavailable local altimeter readings.
Always check the notes for special instructions, adjustments, or requirements like Radar Services Required or what to do when a fix can not be identified.


text:
What makes wing strakes desirable?

Some planes have this large protruding part at the wing root. If my understanding is correct, it's a wing strake and it improves lift at high aoa via vortex induced lift on the rest of the wing. I have 3 questions.
1:If that's true, what are it's downsides?
2:This seems like a desirable thing, why don't more planes have it?
3: What does it have to do with canards? Every single canard aircraft has them, Rutan planes, Velocity planes, basically every canard plane I've ever known has them, and no conventional plane I can think of has them. What's the connection here?

"and no conventional plane I can think of has them" F-5, F-15 (to some extent), F-16, F-18, T-50, SU-27 (including the non-canards versions) just to name a few

I think he means "normal airplanes", not exotics.

@JohnK define "normal airplanes". Also, I would not call several of the most popular fighters in the world "exotic"

"Normal" as in not powered by an engine that puts out more thrust than the thing weighs yadda yadda yadda.  They are exotic from a propeller driven straight wing perspective.  You know, the supersonic deal and all that.  He was providing a propeller driven homebuilt as the basis of his question.

I was meant to say regular general aviation or commercial aircraft. For me, fighter planes are indeed exotic, as I (and probably most other people) only see them rarely at airshows. I'd consider SSTs exotc as well.

The strakes (or root extensions) only have a secondary purpose aerodynamically. Mainly, they solved a particularly vexing problem on the Rutan canard designs (Vari-eze and its offspring), where the wings were swept and the operating C of G  was forward of the wings due to the canard configuration; where to put the consumable load, the fuel, where it can be as close as possible to the C of G.
You couldn't put fuel in the main wings farther aft, because depletion of fuel moved the C of G forward too much, and the Vari-eze already suffered from C of G and main gear placement issues due to the pilot forward, forcing Rutan to use the kneeling camel nose gear arrangement to avoid a tail skid to hold the back end up when empty (a really big problem when there's no tail to speak of, just a propeller).
There was no room in the fuselage for a tank of significant size and still have space for any baggage.  That left the sides of the fuselage ahead of the wing leading edge. He could have just put blisters there for fuel tanks, but that makes drag with no other benefit. The root extensions were an obvious and much more elegant solution.  Minimal drag increase, and a beneficial aerodynamic effect at high AOA, and plenty of fuel volume that changes the balance very little when depleted.
The presence of the strakes moves the mean area of the wing forward somewhat, but minor adjustments to the configuration take care of that.
"Regular" airplanes can put fuel in the straight wings and don't need strakes of that size and if they are used, there are strictly to improve stall behaviour. Some homebuilt owners, like the Thorp T-18 guys, have experimented with small leading edge strakes to improve high AOA behaviour, but these were only about the size of your hand and are just being used as vortex generators (can't find a pic of one so far).

It's pretty common in supersonic aircraft and most have them (Shuttle, Concorde, etc). It brings the center of pressure forward on the plane, and also shifts the lift forward, enabling you to put the wings on the back. Planes with canards often have the winds further back than a traditional plane (velocity, etc) because the canards are in the spot the wings should be. Take a look at the wings on the Concorde below. It's probably not as common as it requires quicker take-off speeds.

"...because the canards are in the spot the wings should be." -- is that really quite what you meant to say?